High bypass ratio turbofan



y' 1968 s'. H. DECHER ,ET AL 3,390,527

HIGH BYPASS RATIO TURBOFAN Original Filed Fab. 26, 1965 3 Sheets-Sheet 1INVENTORS. SlEGFRiED H. DECHER DALE H. RAUCH July 2, 1968 SQH. DECHER Em3,390,527

HIGH BYPASS RATIO TURBOFAN Original Filed Feb. 26, 1965 5 Sheets-Sheet 2INVENTORS. IEGFRIED H. DECHER DALE H. RAUCH ATTORNEYS.

July 2, 1968 s. H. DECHER ET AL 3,390,527

HIGH BYPASS RATIO TURBOFAN Original Filed Feb. 26, 1965 o Sheets-Sheet 3INVENTORS'. SIEGFRIED H. DECHER BY DALE H. RAUCl-j United States Patent3,390,527 HIGH BYPASS RATIO TURBOFAN Siegfried H. Decher, Trumbull, andDale H. Ranch, West Haven, Conn., assignors to Avco Corporation,Stratford, Conn., a corporation of Delaware Continuation of applicationSer. No. 435,515, Feb. 26, 1965. This application July 19, 1967, Ser.No. 657,463 2 Claims. (Cl. 60226) ABSTRACT OF THE DISCLOSURE A highbypass ratio turbofan having a forward fan assembly divided into innerprecompression and outer bypass annular regions, both regions locatedradially outboard of the compressor blading, each uniquely designed fora pressure rise required by aerodynamic conditions, with wall structureextending radially inward and rearward of the inner precompressingportion forming a stationary annular confining passage from the innerprecompressionportion rearward and radially inward into the compressor.

This application is a continuation of our application Ser. No. 435,515,filed Feb. 26, 1965.

This invention relates to a bypass engine and fan assemly, sometimesalso referred to as a turbofan.

It is a principal object to provide a combination of structure whichwill satisfactorily solve certain aerodynamic and stress problemspresented in the use of a relatively high bypass ratio fan in the lowersubsonic flight range and by such structure to obtain performance gainsappreciably higher than in the more commonly used fanjet engines.

Another important feature and object is the use in combination withother elements of a front fan with inner and outer annular portions witha rota-ting integral dividing wall therebetween positioned to serve asthe initial air inlet forward of the engine with the inner fan portioneffecting precompression and divided from the outer bypass region at thewall division in the fan and feeding air through a rearwardly andinwardly extending annular channel to the engine compressor with saidinner supercharging or so-called precompressing portion of the fanlocated outboard and substantially forward of the initial stages of theengine compressor blading.

It is a further feature and object to provide in combination with theaforesaid arrangement a reduction gearing between the power turbinewheel and the fan, such that the relatively higher speed of rotation ofthe power turbine may be obtained, with optimum selection and freedom ofdesign of the fan speed and the turbine speed and number of turbinewheels which may be accomplished by variation in gear ratio.

On consideration of the requirements necessary for the high bypass ratiofan, as compared to the prior art engines using lower bypass ratio wherea fan blade, without dividing wall, was designed to deliver both thebypass air and the air to the engine compressor, it has been found thatwith the larger air flow and lower r.p.rn. of the high bypass ratio fanthat a fan designed to accomplish the dual purposes in this range isdifficult of accomplishment. Particularly, it is not practical to usethe basic design principles of todays low bypass engines by adapting,for example, the fan blade to the larger airflow and smaller fanpressure ratio required. One important reason concerns the fan itself.The much larger airflow demands basically lower rotational speeds in thefan section as compared to the engine compressor. For

high bypass ratios, it becomes therefore impractical to use the designprinciple of certain front fan engines where the fan stages are anintegral part of the compressor.

3,39%,527 Patented July 2, 1968 In any case, a long fan blade of a highbypass fan presents a special aerodynamic problem, if one wants to avoidtoo low root speeds (causing excessively high pres sure coefficients andhigher camber at the root section) or too high tip speeds which willresult in high relative Mach number and loss in efficiency near the tip.

Another reason that low bypass fan principles are not applicableinvolves the fan driving turbine. It is evident that the work outputrequired from the fan driving turbine increases with bypass ratio.Present day aft fan engines, with the turbine section located inside thefan section, are limited with regard to an incerase of work output perstage with increasing bypass ratio. The relatively low rotational speedimposed by the high bypass fan would call for a relatively large numberof turbine stages in case of the high bypass fan engine if gearreduction is not to be used.

It is an important concept of this invention that the aerodynamic fanproblems, mentioned above, can be solved by moving the fan blade rootsection outboard as well as forward of the main compressor. In the caseof the front fan, here disclosed, this requires an annular ducting whichfeeds the engine airflow from the hub portion of the fan blade inwardand rearward toward the front flange of the engine compressor, withsufficient length of passage being afforded to allow for a free streamdiffusion to take place from the fan blading to the inlet of thecompressor. From a mechanical standpoint, moving the fan outboard of theengine compressor makes it possible to bring the blade root stresseswithin acceptable limits. It is a preferred form of this invention touse relatively high-speed power turbine components with one or twostages driving the relatively slow-running fan through a reduction gearfor large bypass ratios (above 4), and with this arrangement the gearratio is in the order of 2.5.

An important distinction over prior art devices results from thecombination of structure afforded by this invention in that certaindesirable advantageous results are obtained without the use of guidevanes ahead of the fan, with their evident disadvantages includingdifliculty of assembly as well as the possibility of icing. This comesabout wherein certain prior art devices use such inlet guide vanes toassist in overcoming the difficulties encountered when the superchargingportion of the fan is left at substantially the same radius as theinitial stages of the engine compressor blading. Under such conditionsthe use of inlet guide vanes ahead of the fan will give the necessaryleverage to run the root speed of the fan at the proper value to solveproblems of high diffusion and to reduce the Mach number at the bladetips by designing such inlet vanes for pro-whirl in the bypass region.It is therefore an important advantage to the herein disclosed structurethat both of these problems are brought Within acceptable limits byestablishing the supercharging region at an annular region outboard ofthe compressor whereby relative speeds are such that a satisfactorycompromise is possible in both inner and outer regions for both stressand aerodynamic considerations and without the objectionable inlet guidevanes.

The above and other objects of the invention will appear more fully fromthe following more detailed description of an illustrative structure andby reference to the accompanying drawings forming a part hereof andwherein:

FIGURE 1 is a schematic cross section showing a gas turbine engine withfront bypass fan rotated by the power turbine and showing location andarrangement of the bypass and compressor supercharging regions of thefront fan and dividing wall therebetween;

FIGURE 2 is a cross section through an engine and fan assembly showingthe upper half of the fan and engine,

including gearing and location of the several elements of thecombination of structure;

FIGURE 2A is a cross section of the front fairing corresponding to andcompleting the showing in FIGURE 2;

FIGURE 3 is a view of the front of the fan assembly partly broken awayand partly in section and taken on the line 3-3 of FIGURE 2.

Referring to the drawings and particularly to the schematic showing ofFIGURE 1 and to the cross section in FIGURE 2, a gas turbine engine ofthe general type shown in US. Patent 3,088,278 is provided with a gasproducer turbine 12, driving a compressor .14 (so-called gas producercompressor), power turbine 16, combustion chamber 18, and air inlet 20.The power turbine 16 rotates a separate power shaft 22 forwardlyextending, and such power shaft 22 is co-axial with and rotatesindependent of, but inside, the gas producer turbine shaft 25. Airenters the inlet and is compressed by the compressor 14 and flows to thecombustor 18. The mixing and combustion of fuel takes place in thecombustor 18 and the hot gas flows into the turbine wheels 12 and 16.This engine has added thereto a front fan assembly 23 and gear reduction24, the structure of which, in combination with certain elements of theengine 10, constitutes the invention here disclosed. The fan assembly 23and gear reduction 24 are made up of an annular fan housing 26 withinwhich is supported a gear train 28 actuated by the shaft 22. This shaft22, which is the power shaft from the power turbine 16, rotates the fanassembly 23 through the gear reduction 24. The fan assembly has an outerbypass fan portion 32 and an inner supercharging fan portion 34. The fanassembly is made up of a plurality of radially extending fan bladessupported from the center shaft 36 and forming two annular regions, onethe outer fan bypass region 32 and the inner fan supercharging portion34 divided by a forwardly projecting wall 38 annularly extending andsupported on the fan assembly 23. The outer housing 26 supports statorblades 4!} rearward of the fan in the outer bypass channel and statorblades 42 in the inner supercharging channel, also rearward of the fan.The outer bypass channel 44 extends rearwardly from the outer bypassportion of the housing 26 into the fan exhaust nozzle 46. The entireexhaust nozzle and the forward housing rearward of the outer bypassportion 32 of the fan 23 is an annular air passage. The air channel 48,rearward of the precompressor fan portion 34, extends in an annular formrearward and inward to the inlet 2% of the air compressor 14.

The fan assembly 23 carries a forwardly extending fairing 50 having adiameter such as to define the inner limit of the supercharging portion34 of the fan assembly, the dividing wall 38 forming the outer dividingportion and outer limit of the inner supercharging fan portion 34 whichon rotation of the fan assembly feeds air into the channel 48 and intothe inlet 20 of the main compressor 14.

The provision of the forwardly extending dividing wall 38 between theinner supercharging portion 34 and the outer bypass portion 32 of thefan assembly 23 affords a positive mechanical wall between thesupercharging and bypass fan regions. In certain prior art devices thedivision in flow between the bypass channel and the inlet to thecompressor is accomplished behind the fan with no forwardly extendingwall, whereas in this disclosure the division begins forward of the fanintegral with and rotating with the fan. The flow division continues ina nonrotating structure 39 behind the fan. (See FIGURE 2.)

It is noted that the annular air channel passage 48 feeds the incomingair flow, from the supercharging portion 34 of the fan, and extendsrearward and inward toward the main compressor 14. Therefore, with theuse of a gear reduction between the turbine shaft 22 and the fan 23 therelative location of the supercharging portion 34 of the fansubstantially outboard of the main compressor 14;,

together with the fairing 58 and the forwardly extending wall 38defining the limits of the supercharging portion of the fan, makes itpossible to make use of a blade form to couple the relatively slowrotating fan to efiiciently deliver the air to the inlet to the maincompressor 14. It is remembered that the relatively slow rotation of thefan 23 is made necessary by the high bypass ratio desired for thisinstallation. Therefore, the fact of positive separation ahead of andintegral with the fan between the outer fan bypass portion and the innersupercharging portion effected by the annular wall 38 is important.

The wall 38 provides a positive mechanical separation. maintaining thepossibility of a positive and etficient change in static pressuregradient fro-m the tip section of the supercharging portion to the rootsection of the bypass portion of each blade. The provision of aseparating wall 38 allows freedom to design the fan portion 32 of theblade and the supercharging portion 34 for different pressure riseswithout any aerodynamic compromise.

The provision of the supercharging or precompressing inner portion ofthe fan at a substantial radius from the center of the fan, in additionto advantages for aerodynamic and stress considerations previouslymentioned, also provides necessary space for the housing of the gearreduction in the central portion of the fan housing assembly, andtherefore such arrangement affords a satisfactory combination forproducing the result desired, particularly advantageous for a fancomponent of high bypass ratio.

The fan assembly 23 is located forward of the engine including the outerbypass region 32 of the fan and the inner supercharging portion 34. Thusthe fan with its flow guiding elements, such as the fairing 5t} and theforwardly extending wall 38, are all rotating parts and are positionedto serve as the initial air inlet forward of the engine. The avoidanceof the use of stationary guide vanes ahead of the fan makes possible adesign of fan blade which will give the desired bypass ratio with speedsof rotation such that the stresses in the rotating blade are kept withinreasonable limits. The reduced speed of rotation is afforded by the gearreduction while still maintaining the high, rotative speed of thedriving turbine, pro vided the inner supercharging portion of the fan islo cated at a substantial radius from the center of rotation of the fan,i.e., outboard of the compressor diameter. All of these features providea combination of structure working together to solve the problems ofstress and aerodynamics in a high bypass fan.

It is Worthy of note that the dividing wall 38 between innersupercharging and outer bypass portions rotates with the blades anddivides the supercharging portion 34 from the bypass portion 32beginning in front of the blade, as distinguished from a division onlyrearward of the blade as in certain prior art devices. The Separatingwall 38 allows the fan portion 32 and the supercharging portion 34 ofthe fan blade to be designed for different pressure rise conditions asthey are in general required for optimum thermodynamic cyclecondition-s. Therefore, the use of a dividing wall beginning forward ofthe fan is an important factor in adapting the high bypass fan engine tothe lower rotational speeds required aerodynamically for the fan toproduce an optimum power plant for subsonic aircraft.

The provision of a gear reduction, such as 28, between the fan and thedriving turbine 16 is a necessary require ment if the relatively highrotational speeds of the turbine are to be maintained. The designproposed where the fan and supercharging blade is moved radiallyoutboard of the main compressor provides the space required for thereduction gearing.

The use of such a gear reduction between the fan driving turbine and thefan assembly, which includes both the bypass fan and the superchargingor precompressing portion, makes possible a complete flexibility indesign between the relative r.p.m. of the fan assembly and the turbine,as well as the relative number of stages of these two elements.

Although the invention has been described by reference to a specificstructure found practical in actual operation, it is intended thatmodifications may be made within the scope of the following claims.

We claim:

1. In a bypass fan and gas turbine engine combination of the typewherein means are provided to cause said fan assembly to operate both tobypass air and to precompress to a gas producer main compressor of saidengine over the entire range of inflight airspeeds including relativelyhigh forward speeds, said means comprising:

an annular outer bypass fan portion of said fan assemannular wallsforming a bypass fan exhaust nozzle rearward of said outer bypass fanportion of said fan assembly;

an inner annular precompressing portion of said fan assembly to feedprecompressed air into said compressor;

an annular forwardly extending wall secured to and rotatable with saidfan and formed to divide said outer bypass fan portion from said innerannular precompressing fan portion and affording positive separationbetween said outer bypass fan portion and inner precompressing fanportion each designed for a different pressure rise required byaerodynamic conditions;

wall structure forming a stationary annular passage extending rearwardand radially inward from said inner precompression portion of said fanassembly into said compressor, with said passage formed to positivelyconfine and to direct flow therethrough;

said annular precompressing portion positioned radial ly outward fromthe extended common axis of rotation of said fan and compressor a radialdistance beginning in a region outboard of the blading diameter of saidcompressor, but only so far outboard as will maintain aerodynamicloading on said fan within acceptable limits when operating inrelatively high or relatively low airspeed ranges in both inner andouter portions;

the spacing of said fan forward of said compressor being in amount suchthat there is afforded a length and configuration of said passageresulting in passage surface velocities which will allow flow into saidcompressor without separation.

2. Structure as in claim 1 in which:

said inner and outer fan portions serve as the initial air inlet forwardof said engine whereby said fan assembly is operable without the use ofinlet guide vanes.

References Cited UNITED STATES PATENTS 943,626 12/1963 Great Britain.

MARTIN P. SCHWADRON, Primary Examiner.

DOUGLAS HART, Assistant Examiner.

